The present invention pertains to the composite structure manufacturing art and, more particularly, to a method of, and apparatus for, fiber lamination.
Composite material technology provides a means to fabricate a high strength, low weight structure. Such technology finds particular application in the aerospace industry.
At present, three principal composite manufacturing methods are practiced. These include tape lay-up, cloth or woven broad good lay-up and filament winding. Each of these manufacturing techniques exhibits serious limitations, as is discussed below.
Tape lay-up composite material manufacturing is comprised of the laying of sections of tape side by side, or in an overlapped relationship, to form the composite structure. This lay-up may be accomplished manually, with mechanical assist, or by fully automatic tape lay-up machines. The fabrication of complex shapes, such as corners, bevels and tapered sections is a tedious process using the tape lay-up method. Often, successive tape sections must be of different widths and/or lengths. A certain amount of overlap must be provided, thereby forming an undesirable "shingle step".
In addition, where the tape is required to form a compound-curve surface, undesirable tape buckling occurs on the concave edge. Further, where changes in direction are required, the tape must be cut and spliced to avoid buckling and strand gapping.
In addition, ply thickness is fixed for a given tape. If changes in either ply thickness or tape width are required, the process must be interrupted to change tape spools.
Cloth good lay-up does not lend itself to mechanization or automation. Further, cutting out contours and laminating these contour shapes is a cost-intensive procedure and often realizes significant material trim losses. In addition, design compromises must be made due to the inability to select any desired fiber angle and the requirement to shingle layer to obtain tapers.
Conventional tape winding manufacturing techniques are deficient in two areas. First, the winding process is generally slow, producing a relatively low productivity rate with limited strand direction capability.
In addition, limitations inherent in the winding process limit manufacturing flexibility. For example, an ideal composite material should have a minimum of gaps between filaments with a maximum ratio of filament area to resin area. Such construction provides the optimum strength-to-weight ratio. Present winding techniques do not realize the optimum ratio. In addition, the filament distribution tends to lack uniformity, such that some filaments are less well supported than others. Further, to make optimum use of the potential of composite materials, the filaments should be consistently and accurately aligned in the direction of the loads being carried, with the quantities of filaments being consistently matched to the load requirements. This requirement can be met relatively easily when the structure and load pattern are geometrically simple. In many applications, however, such as aircraft structure, complicated geometrical shapes must be constructed such that filament winding alignment is necessarily compromised and extra material must be used to meet design requirements. In addition, present winding techniques require overlaps, "shingling", "turn-arounds" and so forth. The net result of these factors is a decrease in strength-to-weight ratio and a limitation in the ability of the manufacturing operation to construct precise contours. Further, the tension winding technique employed in filament winding promotes resin migration and warpage during subsequent autoclave curing.